Gas turbine engines are widely understood to be used in conjunction with generators for turning mechanical shaft power into electrical power. Referring initially to FIG. 1A, a schematic for a simple cycle gas turbine engine 100 commonly used in an electrical generating power plant is depicted. The gas turbine engine 100 comprises a compressor 102 coupled to a turbine 104 by a shaft 106. Air from the compressor 102 is directed to one or more combustors 108 where fuel 110 is added to the air. The fuel and air mixture is ignited to form hot combustion gases which drive the turbine 104, which, in turn, drives the compressor 102. The shaft 106 is also coupled to a generator 112, which produces electric power 114. FIG. 1B shows the corresponding gas turbine performance for thermal efficiency as a function of specific output for both simple cycle efficiency and power output for various gas turbine pressure ratio and firing temperatures. As one skilled in the art understands, the firing temperature of a gas turbine engine regulates and limits the overall operation of the engine and the pressure ratio is directly proportional to the efficiency of the gas turbine. For combined cycle gas turbines, as shown in FIG. 2B, the efficiency of the plant is directly proportional to the firing temperature. In other words, increasing firing temperature increases the output of a simple cycle gas turbine, assuming the mass flow is the held constant, and increases the efficiency of the same gas turbine when operating in combined cycle.
In general, the gas turbine original equipment manufacturers have increased firing temperature by improving the technology of the materials and coatings in the turbine section so hotter gasses can be passed through the turbine while maintaining the capability of the turbine parts.
Referring now to FIG. 2A, a schematic for a combined cycle power plant 200 is depicted and comprises a compressor 202 coupled to a turbine 204 by a shaft 206. Air from the compressor 202 is directed to one or more combustors 208 where fuel 210 is added to the air from the compressor 202. The fuel and air mixture is ignited to form hot combustion gases which power a turbine 204, and drives the compressor 202. The shaft 206 is also coupled to a generator 212, which produces electric power 214. A combined cycle power plant 200 also includes a heat recovery steam generator, or HRSG, 216, which receives hot exhaust from turbine 204 and heats a water source to generate steam 218. A steam turbine 220 is powered with steam from the HRSG 216, with the steam turbine 220 driving a second generator 222 for generating additional electrical power 224. FIG. 2B shows the corresponding gas turbine performance for efficiency as a function of firing temperature for both the combined cycle efficiency and power output. FIGS. 1B and 2B are similar to those disclosed in GE Gas Turbine Performance Characteristics (GER3567) and are included herein for reference purposes.
As one skilled in the art understands, firing temperature is defined as the temperature of the combustion gases just downstream of the first stage turbine nozzle. Due to different terminology used in the field of gas turbine engines, the first stage turbine nozzle may also be referred to as a first stage turbine vane. Referring to FIG. 3, a cross section of a portion of a gas turbine engine is depicted and indicates standard temperature parameters utilized in the gas turbine industry. FIG. 3 is also similar to that disclosed in the GE Gas Turbine Performance Characteristics (GER3567) paper referenced above. As shown in FIG. 3, turbine inlet temperature (TA) is measured upstream of a first stage turbine nozzle 300, as depicted by plane A-A. The firing temperature of the engine (TB) is measured just aft of the first stage turbine nozzle, as depicted by plane B-B.
As discussed above, turbine inlet temperature and turbine firing temperature are critical measures by which gas turbine engine operation is based. These temperature readings are taken upstream and downstream of the first stage turbine nozzle, respectively. As such, it is important for the turbine nozzle metal temperature to be maintained within acceptable material operating limits as control of the gas turbine engine is based off of these temperatures.
Due to the high operating temperature of the turbine nozzle, it is necessary to actively cool the turbine nozzle in order to maintain metal temperatures at an acceptable level. Cooling fluid, such as compressed air, is provided to the turbine nozzle as part of the overall Turbine Cooling and Leakage Air (TCLA), or compressed air bypassing the combustion process and used for cooling. TCLA is typically taken from multiple locations in the compressor, including the discharge plenum of a gas turbine engine, with the amount required for cooling turbine components varying by component and by engine type. However, for a General Electric Frame 7FA engine, approximately 20% of the compressed air generated by the engine compressor is used as TCLA. That is, using 20% of the compressed air for cooling means this air cannot go through the combustion system, or is unfired going through the turbine, thereby translating into lost energy for the engine and contributing to the poor thermal efficiency of the gas turbine engine. For example, the aforementioned gas turbine engine has a thermal efficiency of approximately, which is approximately 37 percent.
FIG. 4, which is similar to that disclosed in GE Gas Turbine Performance Characteristics (GER3567), depicts a typical cooling scheme for a first stage turbine nozzle 400. In such a cooling arrangement, compressed air is supplied to an internal passage of the turbine vane and is often directed through a plurality of passageways within the nozzle, some of which can be serpentine in shape. The air for cooling the first stage turbine nozzle is typically produced by the compressor and is taken from a compressor discharge plenum and therefore is at the exit pressure and temperature of the engine compressor. This first stage nozzle, which sees the highest temperature gases from the combustor, is also supplied with the sources of highest pressure cooling air, from the compressor discharge plenum (CDP). That is, the pressure of the gas path is just a couple of pounds per square inch (psi) less than that of the combustor. Therefore, as one skilled in the art can appreciate, the pressure of the cooling air supplied to the leading edge 402 of the first stage nozzle 400 is just high enough to cause air to flow out a series of holes in the airfoil. Cooling hole spacing and orientation can vary, but one such common style places holes in the leading edge 402 of the nozzle 400, also referred to as a showerhead pattern. Further, taking air from the engine compressor to cool the turbine components reduces the power output from the engine, and thus the amount of mechanical work able to be generated by the turbine.
Referring now to FIG. 5, a cross section view of a portion of a gas turbine engine in accordance with a cooling scheme of the prior art is depicted. The gas turbine engine 500 comprises a compressor 502 providing a flow of compressed air into a discharge plenum 504. Most of the air from the compressor 502 passes through one or more combustors 506, the one or more combustors 506 having a combustor case 508, an end cap 510, a combustion liner 512, a swirler assembly 514, a transition piece 516, and a bracket 518 that holds the transition piece 516 to a portion of a turbine frame, here the first stage vane outer ring 520. Air is received in the combustor 506 and mixed with fuel from one or more fuel nozzles 522 to create hot combustion gases passing through the transition piece 516 and into the turbine. In this embodiment, the first stage vane outer ring 520 is fastened to the compressor discharge plenum (CDP) case 524.
Air is maintained in the compressor discharge plenum by seal 526 between the rotor 528 and an inner casing 530 such that most of the air goes to the combustor 506 or for TCLA. The inner casing 530 has a mechanical interface 532 with the first stage turbine nozzle 531 for providing needed structural axial and torsional support. The inner casing 530 is generally supported within compressor discharge plenum case 524 by ID struts 534 located between adjacent combustors 506. The rotor 528 has bearings 536 that tie the rotor 528 to the casing through struts 534.
The cooling air 541 is supplied to the outer diameter of the first turbine nozzle 531 and passes between the first outer vane ring 520 and the compressor discharge plenum case 524 and enters into holes on the first vane outer ring 543 as the first vane outer ring feeds the vane 531 with compressed air from the compressor discharge plenum 504. In this embodiment of the present invention, the compressed air from the compressor discharge plenum 504 is approximately 750 deg. F. at ISO conditions and base load. Similarly, the inner diameter of the first stage nozzle 542 is supplied with turbine cooling and leakage air (TCLA) 552 from the compressor discharge plenum 504. Both first stage nozzle cooling air 541 and 552 flows through the internal passages 531 of the vane, as disclosed in FIG. 4, providing the necessary cooling to the first stage nozzle 542. Eventually this TCLA joins with the hot combustion gases passing between the first stage nozzles 542 and acts as a coolant to reduce the temperature of the hot gases to which the first stage blade 511 is exposed. On subsequent nozzle and rotor stages, the second stage nozzle is sealed to the rotor with a second stage inner support ring 554 and similarly on the third stage with a third stage inner support ring 553.
The following discussion pertains to a General Electric Frame 7FA gas turbine engine at ISO conditions and base load and is provided merely for illustrative purposes as an acceptable engine with which the present invention can be utilized and is not meant to limit the scope of the invention discussed below. The majority (about 80%) of the compressed air from the compressor passes through the combustion system where fuel is added and the mixture is ignited, raising the temperature of the hot combustion gases to approximately 2700 deg. F. There is typically a two to three pound per square inch (psi) pressure drop as the compressed air goes through the combustor. Therefore, because of this arrangement, there is very little pressure margin to cool the nozzle, especially its leading edge. Typically on an F-class gas turbine engine, approximately 10% of the cooling air is diverted from the combustion process and is used to cool the vane. For example, for the 7FA engine, compressor discharge air at approximately 750 deg. F. and 220 psi is used to cool the first stage nozzle. During the cooling process, this air increases in temperature by approximately 250 deg. F. and is then discharged into the gas path, thereby diluting the hotter (2700 deg. F.) temperature gasses coming from the combustion process, yielding a firing temperature. A typical firing temperature for the 7FA engine is approximately 2450 deg. F. (as taken at plane B-B) and comprises 900 lb/sec of hot combustion gasses at a temperature of approximately 2700 deg. F. from the combustion process and 100 lb/sec of air at approximately 1000 deg. F. from the cooling air for the nozzle. Therefore, this yields a firing temperature of 2540 deg. F. at plane B-B [(2700*900+100*1000)/1000=2540 deg. F.]. The reason for the higher temperature in the calculation (2540 F>2450 F) is because there is also some combustion dilution and cooling air that mixes out and reduces the actual temperature exiting the combustor, therefore, reducing the temperature at plane B-B. To estimate the effective combustion dilution and leakage air which is at compressor exit temperature (750 deg. F.), (2700*900+100*1000+Flow*750)/(1000+Flow)=2450, and when solving for the flow, Flow=5. Therefore, with a compressor inlet flow of approximately 1005 lb./sec, 900 lb./sec goes through the combustion process, and approximately 5 lb./sec leaks and dilutes the combustion process and 100 goes to the first stage nozzle cooling. These numbers do not reflect the fact that in the compressor of the gas turbine, approximately 10% of the 1005 lb./sec going to the turbine inlet is removed before it exits the combustor in order to cool the rotating section and later static sections of the turbine. Therefore, for the example discussed above, all the flow numbers are reduced by 10%, or the combustor flow is approximately 810 lb./sec, the first stage nozzle flow is approximately 90 lb./sec and the combustor dilution and leakage rate is 4.3 lb./sec. As one skilled in the art can appreciate, these numbers are approximate, however, when the leakage and cooling air is mixed in plane B-B, a blended temperature of 2450 deg. F. (firing temperature) results.
An industry standard for determining the cooling benefit achieved through the cooling air is its cooling effectiveness. Cooling effectiveness is understood to be the ratio of the difference between the hot combustion gas temperature and the average metal temperature of the turbine nozzle divided by the difference between the hot combustion gasses and the temperature of the cooling air. As an example, the cooling effectiveness of the first stage turbine vane of the 7FA engine discussed above is approximately 0.59 (the ratio of the temperature difference between the hot combustion gasses (˜2700) and average metal temperature (˜1550) divided by the difference between the hot combustion gasses and cooling air temperature (˜750 F)).
Cooling the highest temperature components, typically the first stage nozzles and first stage blades, is a technology on which every gas turbine engine original equipment manufacturer (OEM) spends significant financial resources. For example, over the last twenty years, large frame gas turbine engines have been improved, but thermal efficiency improvement has risen from about 33% to only about 37%.